1. Field of the Invention
The present invention relates go gas turbine engines, and more specifically to a high pressure ratio compressor with last stage airfoil cooling used in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an aero engine used to power an aircraft or an industrial gas turbine engine used to produce electric power, is a very efficient power plant. The compressed air from a compressor is passed into a combustor where the air is burned with a fuel to produce a hot gas flow, the hot gas flow is then passed through a turbine to drive the rotor shaft connected to the compressor and, in the case of an aero engine produce thrust and/or drive the fan, or in the case of the IGT drive an electric generator to produce the electric power. In both cases, the efficiency of the engine can be increased by passing a higher temperature gas into the turbine.
Modern gas turbine engines have multiple stages in the compressor in order to produce the very high pressure ratios between the outlet of the compressor and the inlet. For example, the Pratt & Whitney F100 aero engine that is used to power the military F15 and F16 fighter aircraft includes 13 stages in the compressor and produces a pressure ratio of 30 to 1 (the outlet pressure is 30 times the inlet pressure). A higher pressure ratio will allow for higher efficiencies for the engine. With the recent improvements in compressor design, a higher number of stages can be used to produce an even higher pressure ratio. Futures aero engines are anticipated to have compressor ratios in the 50s.
However, as the air through the compressor is compressed, the temperature of the compressed air also increases. A multiple stage compressor will generally add 90 degrees F. to the compressed air for each stage. As the number of stages in the compressor grows, the compressor outlet air becomes higher to the point where the last stage airfoils (stator vanes and compressor blades) become so hot that the airfoils can be damaged from the high thermal load. Thus, there is a need in the prior art for a multiple stage compressor with a very high pressure ratio to have cooling of the last stage airfoils in the compressor in order to withstand the higher temperatures.